Turbine blade for a gas turbine engine

ABSTRACT

A turbine blade is manufactured as a single-crystal casting from a metal alloy, without a solution heat treatment step.

This invention relates to a turbine blade for a gas turbine engine, theuse of such a turbine blade in a gas turbine engine, and to a method ofmanufacturing a turbine blade.

Turbine blades in gas turbine engines operate at the limits of theirmaterial properties. They may be exposed to temperatures in excess of2000° C. and are subjected to severe stress, both from gas flow past theblades and from centrifugal forces.

It is known to form turbine blades as single-crystal castings fromspecialized metal alloys, thereby providing high strength required toavoid failure under operational loads. The alloys used tend to benickel/aluminum alloys, with various other components selected toenhance the properties of the alloy. Typical alloys used formanufacturing turbine blades for gas turbine engines are disclosed, byway of example, in U.S. Pat. No. 4,222,794, U.S. Pat. No. 4,582,548,U.S. Pat. No. 4,643,782 and U.S. Pat. No. 5,540,790. Some of the alloysdisclosed in these patent specifications are commercially available, forexample under the designations CMSX-3, CMSX-4 (available from CannonMuskegon Corporation of Muskegon, Mich., USA) and PW 1484.

It has been considered essential for turbine blades manufactured assingle-crystal castings to be heat treated before use to relieveresidual stresses, thereby optimizing the mechanical properties of thealloy. Early single crystal alloys were not heat treatable because thetemperature at which the necessary strengthening changes occurred wasabove the melting point of the material. Hence such alloys were not usedto produce turbine blades because their poor microstructure inherentlymeant their mechanical integrity was insufficient for such applications.Improved single crystal materials are now available which enablesolution heat treatment, thus delivering optimal mechanical properties.

Residual stresses in single-crystal castings arise as a result ofdifferential contraction of different parts of the casting as it cools.Solution heat treatment relieves these stresses but a disadvantage isthat re-crystallization of the material may occur, which will weaken thestructure.

Re-crystallization can be minimized by appropriate design of the turbineblade. In particular, it appears that re-crystallization is inhibited ifinternal webs within the turbine blade extend perpendicular to, or closeto perpendicular to, the external walls of the turbine blade, if thewebs are relatively thick, and if the spacing between adjacent coolingholes is relatively large. However, a turbine blade designed withinthese constraints may not have optimum performance. For example, thickerwebs increase the weight of the blade, while the angles of the websrelative to the outer walls of the blade and the spacing of coolingholes can affect the cooling efficiency of the blade, in terms of thequantity of cooling air required to maintain a desired temperature.

Nevertheless, it has until now been believed that solution heattreatment of single-crystal turbine blades provides the only route bywhich an acceptable operational life can be achieved, and solution heattreatment has therefore been regarded as an essential step in themanufacture of such turbine blades.

According to one aspect of the present invention, there is provided afinished turbine blade for a gas turbine engine, comprising asingle-crystal casting of a metal alloy having a solvus temperaturewhich is less than its incipient melting point, characterised in thatthe turbine blade has not been subjected to solution heat treatmentafter casting.

Another aspect of the present invention provides the use in a gasturbine engine of a turbine blade comprising a single-crystal casting ofa metal alloy having a solvus temperature which is less than itsincipient melting point, characterised in that the turbine blade has notbeen subjected to solution heat treatment after casting.

A further aspect of the present invention provides a gas turbine engine,characterised in that the engine includes a turbine blade comprising asingle-crystal casting of metal alloy having a solvus temperature whichis less than its incipient melting point, characterised in that theturbine blade has not been subjected to solution heat treatment aftercasting.

Thus, while it was previously considered to be essential to heat treat aturbine blade formed as a single crystal casting in order to achievedesired physical properties to ensure the required operational life ofthe component, the present invention arises from the realization thatthe additional design flexibility which arises if solution heattreatment is avoided can compensate for any resulting deficiencies inthe physical properties of the material which would otherwise lead to areduced operational life.

The metal alloy from which the turbine blade is made is preferably anickel-based alloy, such as a nickel aluminum alloy, including SRR99,CMSX-3, CMSX-4 and PWA 1484. The alloy may contain other alloyingcomponents, such as hafnium, rhenium, titanium, chromium or gallium. Thesolvus temperature of the metal alloy should be less than the incipientmelting point of the alloy.

The turbine blade may be provided with internal cooling passages in theform of cavities extending through the blade. By virtue of the designfreedom which results from the omission of any solution heat treatmentstep, internal walls within the turbine blade, which separate adjacentcavities from one another, may be thinner than in a turbine blade whichis subjected to a solution heat treatment step. For example, thethickness ratio between internal walls which separate adjacent cavitiesfrom one another and external walls which separate the cavities from theexterior of the turbine blade, may be less than 1.5:1 and preferablyless than 1.25:1. Also, the angle at which an internal wall meets theexternal wall may be smaller than in a turbine blade which has beensubjected to a solution heat treatment step. For example, an internalwall may meet an external wall at an angle less than 60° and possiblyless than 50° or 45°.

The internal walls may be provided with through holes, for example forthe passage of cooling air, and these holes may be more closely spacedthan in a turbine blade that has been subjected to solution heattreatment. For example, the centreline spacing between adjacent holesmay be less than 6 times the hole diameter, or even less than 5 times or4 times the hole diameter.

Another aspect of the present invention provides a method ofmanufacturing a turbine blade for use in a gas turbine engine, themethod comprising casting the turbine blade as a single crystal from ametal alloy, without a subsequent solution heat treatment step, saidmetal alloy having a solvus temperature less than its incipient meltingpoint.

For a better understanding of the present invention, and to show moreclearly how it may be carried into effect, reference will now be made,by way of example, to the accompanying drawings, in which:

FIG. 1 (PRIOR ART) is a sectional view of a turbine blade for a gasturbine engine, the manufacture of which includes a solution heattreatment step;

FIG. 2 corresponds to FIG. 1, but shows a turbine blade configurationsuitable for a turbine blade which is not subjected to solution heattreatment after casting;

FIG. 3 (PRIOR ART) is a sectional view on the line A-A in FIG. 1; and

FIG. 4 is a sectional view taken on the line B-B in FIG. 2.

The turbine blade shown in FIG. 1 is hollow, comprising a plurality ofcavities 2 through which, in use, cooling air may flow in order to coolthe turbine blade. Thus, the cavities 2 may be connected at one or bothends to cooling air supply chambers or ducts situated externally of theturbine blade itself.

The cavities 2 are separated from each other by internal walls 4.Cooling holes 6 are provided which extend through the walls 4 to allowthe flow of cooling air between adjacent cavities 2. The cavities 2 arealso bounded by external walls 6 which separate the cavities 2 from theoutside of the turbine blade.

The turbine blade shown in FIG. 1 is cast as a single crystal from asuitable single-crystal alloy, for example CMSX-3, CMSX-4 or PW1484. Ithas been considered to be essential for the turbine blade, aftercasting, to be subjected to a solution heat treatment process in orderto relieve residual stresses in the cast turbine blade. Solution heattreatment processes have been required in order to enhance the strengthof the cast turbine blade, and in particular to enhance the fatiguestrength and creep strength. Consequently, alloys for use in themanufacture of single-crystal turbine blades have been formulated so asto have a solvus temperature (ie the temperature at which the necessarystrengthening changes will occur) which is below the incipient meltingpoint of the alloy. Consequently, the beneficial changes achieved bysolution heat treatment occur before the cast turbine blade begins tomelt.

However, the solution heat treatment process is known to causerecrystallization of the alloy, which weakens the structure in theregions at which recrystallization occurs. The configuration shown inFIG. 1 is designed to minimize such recrystallization. Thus, theinternal walls 4 are relatively thick, the angles at which they meet theexternal walls 6 are relatively large, and the spacing between adjacentcooling holes 8 is also relatively large.

By way of example, in the configuration shown in FIG. 1, the internalwalls 4 have a thickness T_(i) which is significantly larger than thethickness T_(e) of the external walls 6. For example, the ratioT_(i)/T_(e) is greater than 1:5:1. In the embodiment shown, it isapproximately 3:1.

Also, the angles at which the internal walls 4 meet the external walls6, as measured between the general central axis of the internal wall 4in the section shown in FIG. 1 and the tangent to the external surfaceof the external wall 6 at the point of intersection with the centralaxis of the internal wall 4, as designated by way of example by α inFIG. 1, is preferably as close to 90° as possible. Although this cannotbe achieved if the internal walls 4 are to be generally straight, owingto the curved and tapering nature of the turbine blade, it is preferablefor the angle α to be no less than 60°.

As shown in FIG. 3, the cooling holes 8 of the turbine bladeconfiguration of FIG. 1 are disposed in a single line at a centrelinespacing S which is relatively large. Expressed as a multiple of thediameter D of each cooling hole, the spacing S is preferably at least 5times the diameter D and, in the embodiment shown in FIG. 3, isapproximately 7 times the diameter D.

The configuration shown in FIGS. 1 and 3, established with a view to aminimizing recrystallization of the material of the turbine blade duringsolution heat treatment, has disadvantages. In particular, therelatively thick internal walls 4 increase the volume of alloy in theturbine blade, and consequently also increase the weight of the turbineblade. Also, the relatively thick internal walls 4 reduce the sizes ofthe cavities 2, so reducing the flow passage size for cooling air.Cooling air flow is also restricted by the relative spacing S betweenadjacent cooling holes 8, since this restricts the number of cooling 8that can be provided. The need for a relatively large angle α makes itimpossible to angle the internal walls 4 relatively to the externalwalls 6 by narrow angles, which could allow the cooling holes 8 to bedirected at the external walls 6, so as provide impingement cooling.

These disadvantages are overcome in the turbine blade configurationshown in FIGS. 2 and 4, in which the features of the turbine blade aredesignated by the same reference numbers as in FIGS. 1 and 3.

In the configuration shown in FIGS. 2 and 4, the internal walls 4 aresignificantly thinner than those of the configuration shown in FIG. 1.In the case of some of the walls 4, the thickness T_(i) is comparable tothe thickness T_(e) of the external walls 6, so that the ratioT_(i)/T_(e) is close to 1. Generally, it is preferred for the ratioT_(i)/T_(e) to be no greater than 1.5:1, or more preferably, no greaterthan 1.25:1.

Furthermore, some of the internal walls 4 meet the external walls 6 atangles α significantly less than the corresponding angles α of theconfiguration shown in FIG. 1. For example, the angle α for at leastsome of the internal walls 4 in the configuration of FIG. 2 may be lessthan 60° or even less than 50°, and in some cases may be 45° or smaller.

As shown in FIG. 4, the spacing between adjacent cooling holes 8 issignificantly smaller than that of the configuration of FIG. 3. Thus,the spacing S may be less than 6 times the diameter D of each coolinghole, or even less than 5 times the diameter D. In the embodiment shownin FIG. 4, the spacing S is only about 3 times the diameter D. Thisenables the cooling holes to be arranged in two rows, which would not bepossible in the configuration shown in FIG. 3 if the spacing S is to besufficiently large to avoid recrystallization of the alloy duringsolution heat treatment.

As a result of the greater design freedom applicable to theconfiguration shown in FIGS. 2 and 4, the reduced thickness T_(i) of theinternal walls 4 results in a weight reduction and an increase in theflow capability of the cavities 2. The ability to use a relatively acuteangle α between the internal walls 4 and the external walls 6 means thatthe cooling holes 8 can be oriented so that they can direct cooling aironto a nearby external wall 6, so providing impingement cooling. Also,the ability to decrease the spacing S between cooling holes, means thatthe number of holes 8 can be increased, which not only increases theflow of cooling air, but also increases the surface area available forthe transfer of heat from the alloy of the turbine blade to the coolingair.

Thus, although the process of manufacturing the turbine blade shown inFIGS. 2 and 4 without solution heat treatment means that, in somerespects, the strength of the turbine blade may not be fully optimized,the advantages arising from weight reduction, increased cooling air flowcapability and effective orientation of the cooling holes 8 means thatthe loss of fatigue and creep strength are outweighed. As a result,contrary to expectations, a turbine blade manufactured to theconfiguration shown in FIGS. 2 and 4 can have an operational lifesimilar to, or exceeding, that of a turbine blade having theconfiguration of FIGS. 1 and 3, despite the fact that the turbine bladeof FIGS. 2 and 4 is manufactured without a solution heat treatment stepas used in the turbine blade of FIGS. 1 and 3.

The omission of the solution heat treatment step has the additionaladvantage that the overall manufacturing time and cost is reduced.Furthermore, the rate of rejection of turbine blades manufactured assingle-crystal cast alloy components can be reduced, since many turbineblades are rejected largely as a result of an unacceptable degree ofrecrystallization during the solution heat treatment process.

1-6. (canceled)
 7. A gas turbine engine, characterised in that theengine includes a turbine blade comprising a single-crystal casting ofmetal alloy having a solvus temperature which is less than its incipientmelting point, characterised in that the turbine blade has not beensubjected to solution heat treatment after casting.
 8. A gas turbineengine as claimed in claim 7 wherein the metal alloy is a Ni ALX alloyin which X is one or more of hafnium, rhenium, titanium, chromium orgallium.
 9. A gas turbine engine as claimed in claim 7 wherein theturbine blade is provided with cavities which are separated fromadjacent cavities by internal walls and which are separated from theexterior of the turbine blade by external walls.
 10. A gas turbineengine as claimed in claim 9 wherein the thickness ratio Ti/Te of theinternal and external walls is not greater than 1.5:1.
 11. A gas turbineengine as claimed in claim 9 wherein at least one of the internal wallsis provided with through holes, at least some of which are spaced apartby a spacing(s) which is not greater than 6 times the transversedimension of the holes.
 12. A gas turbine engine as claimed in claim 9wherein at least one of the internal walls meets an adjacent externalwall at an angle α which is not greater than 60°.
 13. A finished turbineblade for a gas turbine engine, comprising a single-crystal casting of ametal alloy having a solvus temperature which is less than its incipientmelting point, characterised in that the turbine blade has not beensubjected to solution heat treatment after casting.
 14. A turbine bladeas claimed in claim 13 characterised in that the metal alloy is a NiAlXalloy in which X is one or more of hafnium, rhenium, titanium, chromiumor gallium.
 15. A turbine blade as claimed in claim 13 characterised inthat the turbine blade is provided with cavities which are separatedfrom adjacent cavities by internal walls and which are separated fromthe exterior of the turbine blade by external walls.
 16. A turbine bladeas claimed in claim 15 characterised in that the thickness of ratioTi/Te of the internal and external walls is not greater than 1.5:1. 17.A turbine blade as claimed in claim 15 characterised in that at leastone of the internal walls is provided with through holes, at least someof which are spaced apart by a spacing(s) which is not greater than 6times the transverse dimension of the holes.
 18. A turbine blade asclaimed in claim 15 characterised in that at least one of the internalwalls meets an adjacent external wall at an angle α which is not greaterthan 60°.
 19. A method of manufacturing a turbine blade for use in a gasturbine engine, the method comprising casting the turbine blade as asingle crystal from a metal alloy, without a subsequent solution heattreatment step, said metal alloy having a solvus temperature less thanits incipient melting point.